Pulse rocket engine

ABSTRACT

A liquid propellant rocket engine comprising a chamber body connected to a converging/diverging thrust nozzle and to at least one lightly pressurized propellant feed duct, said body including a body bore in which an injector-forming piston can move longitudinally between a rest first position and an operating second position, the piston comprising a piston head and a piston rod, and subdividing the chamber body bore into a combustion chamber in front of the piston, and at least one injection chamber surrounding the piston rod, the chamber body further including a flow section that is closed in said rest position to prevent any propellant being fed from the injection chamber(s) and that is opened when the piston moves from said rest position to said operating position in order to allow the injection chamber(s) to be filled via said flow section directly from the feed duct(s), the propellant(s) being injected into the combustion chamber via injection channels formed between the combustion chamber and the injection chamber(s), displacement of the piston being a function, in particular, of the pressures in the feed duct(s), in the injection chamber(s), and in the combustion chamber.

FIELD OF THE INVENTION

The present invention relates to a rocket engine capable of operatingintermittently with very high combustion pressure and designed moreparticularly for miniaturized thrust systems such a those that may beused in satellites or in interceptor vehicles.

PRIOR ART

It is possible to generate pulses of thrust from a conventional rocketengine architecture that normally produces thrust continuously by actingon the control sequence applied to the injection valve of the engine.Thus, by successively exciting said valve and not exciting said valve,it is possible to cause successive pulses of thrust to be deliveredperiodically, thus enabling the engine to operate intermittently.However, with such a conventional architecture, it is practicallyimpossible to generate pulses of a fixed value that is reproducible andefficient, in particular because of the large amounts of "dead" volume,i.e. volume that exists between the injection valve and the injector andthat is necessarily included in that type of rocket engine.

Thus, special rocket engine architectures have been developed that arespecifically adapted to pulse operation. In application FR 2 446 385filed in 1979 by Rockwell International Corporation, the rocket enginedescribed includes a differential type piston that subdivides thecombustion chamber of a loading chamber into which one or two liquidpropellants can be introduced from storage tanks under the action ofcentrifugal force. The piston is displaced by the action of a controldevice which, in the example shown, is present in the form of a tank ofgas under pressure that feeds the loading chamber via a solenoid valve.In application FR 1 565 785, filed in 1968 by Bolkow GmbH, theinstallation shown for feeding propellant comprises a cylinder, adifferential piston, two metering chambers designed to receive theliquid propellants, and an expansion compartment connected to thecombustion chamber by a central pipe along which the combustion gasesflow under pressure. The metering chambers naturally refill after eachstage of operation from propellant storage tanks that are under lowpressure, and non-return valves are connected at the inlets to saidchambers to prevent the propellants flowing back into the tanks.Similarly, high pressure valves are inserted at the outlets from themetering chambers to enable or prevent propellant injection into thecombustion chamber. Combustion is started by operating the controllevers of the valves that release the propellants into the combustionchamber.

Compared with conventional architectures, the above-described deviceshave the advantage of being suitable for being dimensioned as a functionof a single maximum determined operating time, thereby making itpossible to use a structure that is much lighter in weight than are thestructures of prior devices. Unfortunately, they still suffer fromseveral drawbacks. In particular, combustion is started (the engine isput into operation) by means of special control mechanisms (a tank underpressure and a solenoid valve, high pressure valves), therebyconsiderably increasing the overall mass of the device (which is also ofgreat importance in space applications) and such means are complex tocontrol. In addition, the presence of non-return elements acting on thepropellants or the combustion gases is essential for implementing thatrocket engine, and if one of those elements should break or cease tooperate there is a danger that the entire operation of the engine willbe interrupted, thereby causing the corresponding mission to be lost. Inaddition, it is difficult to dimension such non-return elements, and itis generally necessary to provide a plurality thereof, as specified inabove-mentioned patent application FR 2 446 385.

OBJECT AND BRIEF DEFINITION OF THE INVENTION

An object of the present invention is to provide a pulse rocket enginewhich mitigates the drawbacks of prior art devices by proposing a novelarchitecture which is more compact and cheaper, and which alsoguarantees greater security for thrust systems that use said engine,while nevertheless being entirely compatible for use with conventionalthrust systems in spite of the very high combustion pressure that thenovel architecture implies.

This object is achieved by a liquid propellant rocket engine comprisinga chamber body connected to a converging/diverging thrust nozzle and toat least one lightly pressurized propellant feed duct, said bodyincluding a body bore in which an injector-forming piston can movelongitudinally between a rest first position and an operating secondposition, the piston comprising a piston head and a piston rod, andsubdividing the chamber body bore into a combustion chamber in front ofthe piston, and at least one injection chamber surrounding the pistonrod, wherein the chamber body further includes a flow section that isclosed in said rest position to prevent any propellant being fed fromthe injection chamber(s) and that is opened when the piston moves fromsaid rest position to said operating position in order to allow theinjection chamber(s) to be filled via said flow section directly fromthe feed duct(s), the propellant(s) being injected into the combustionchamber via injection channels formed between the combustion chamber andthe injection chamber(s), displacement of the piston being a function,in particular, of the pressures in the feed duct(s), in the injectionchamber(s), and in the combustion chamber.

This specific structure avoids the drawbacks of the prior art since itbecomes possible to generate pulses of thrust automatically one afteranother without resorting to external control means for starting eachcycle, and without resorting to the use of high pressure valves. Inaddition, the architecture is simplified since the flow passage forpropellant is closed during return of the injector, thereby isolatingthe injection chamber(s) from the feed duct(s), and thus avoiding anyneed for conventional non-return valves or other, analogous components.

In a particular embodiment, the body of the combustion chamber asextended by a thrust nozzle is secured to the head of the piston to forma moving assembly capable of being displaced longitudinally in theengine body, from the rest position in which the flow section is closedto the operating position in which said section is opened to allow theengine to be fed with propellant.

The flow section is preferably implemented in the form of a recessformed at the inlet to the injection chamber(s) in the bore of the bodyalong which the piston rod slides.

In the embodiment under consideration, the channels for injectingpropellant into the combustion chamber can also be constituted by theclearance that exists between the bore of the chamber body and theperipheral surface of the piston head.

In a variant embodiment that enables engine operation to be optimized,the piston head includes a forwardly projecting portion on its frontface that faces towards the combustion chamber, which portion can closethe throat of the thrust nozzle in full or in part on displacement fromthe rest position to the operating position.

The rocket engine of the invention also includes a downstream stop thatis preferably constituted by a shoulder on the thrust nozzle, and thatis designed to limit displacement of the piston in said operatingposition.

When the engine is fed with two propellants, the piston rod is annularin shape and defines both a central injection chamber and an annularinjection chamber, said piston rod including a wall that projectsrearwards and that slides in a second annular cavity which is in coaxialalignment with a first annular cavity in which the piston rod slides, insuch a manner that said wall defines two concentric cavities for feedingthe two propellants from their respective feed ducts. This wall may bereplaced by at least one resilient bellows that is fixed, preferably bywelding, to an upstream end of the chamber body.

Advantageously, in this two-propellant version, the piston rod slidesboth externally in a bore of the chamber body and internally on aplunger-forming part that is fixed to the upstream end of said body byresilient fixing means, in such a manner that the connection between thepiston and the body which is made slightly flexible in this way limitsany risk of the piston jamming during displacement, while still ensuringgood sealing.

When only one propellant is fed to the engine, the combustion ignitionmeans advantageously comprise a catalytic bed disposed on the front faceof the piston head that faces towards the combustion chamber. Inaddition, return means may be provided for holding the piston in therest, first position, in particular when there is no propellant feed.

To stop its operation, the rocket engine of the invention may includelocking means for locking the piston in the rest, first position, saidlocking means may be constituted either by an electromechanical unitcomprising an electromagnet which, when excited, causes a locking fingerto move into an orifice formed in the piston rod or in any other partsecured thereto, or else by an electrohydraulic unit formed by anactuator fed from at least one feed duct via a solenoid valve and onwhich there can slide a cavity formed in the piston rod, in such amanner that said piston rod can be locked in a determined position underthe effect of the pressure forces involved. When at least a portion ofthe piston, preferably the piston rod, is made of a ferromagneticmaterial, and at least the upstream end of the chamber body is made of anon-ferromagnetic material, then the locking means may advantageously beconstituted by an electromagnetic unit which causes displacement of thepiston to be stopped whenever it is excited by control means.

In a particular embodiment, the rocket engine may have non-circularsliding geometry for the piston and it may have a nozzle throat sectionthat is substantially rectangular.

BRIEF DESCRIPTION OF THE FIGS.

Other characteristics and advantages of the present invention appearbetter from the following description, given by way of non-limitingindication, and with reference to the accompanying drawings, in which:

FIG. 1 is a diagrammatic longitudinal view of a rocket engine of thepresent invention fed with a single propellant;

FIGS. 2a and 2b are operating diagrams for a pulse rocket engine of theinvention;

FIGS. 3 and 3a are diagrammatic longitudinal views of rocket engines ofthe invention fed with two propellants;

FIG. 4 is a variant embodiment of the rocket engine of FIG. 3;

FIG. 5 is a theoretical diagram for a variant embodiment of the singlepropellant rocket engine of FIG. 1;

FIG. 6 is a theoretical diagram of another variant embodiment of the twopropellant rocket engine of FIG. 3;

FIGS. 7a and 7b show in highly diagrammatic manner variant embodimentsof certain aspects of the FIG. 6 rocket engine; and

FIG. 8 is a theoretical diagram of yet another variant embodiment of thetwo propellant rocket engine of FIG. 3.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

Reference is made to FIG. 1 which is a diagrammatic longitudinal sectionthrough a pulse rocket engine of the invention fed with a singlepropellant. The rocket engine comprises a circularly symmetrical body 10about a longitudinal axis having a converging/diverging thrust nozzle 12connected to the downstream end thereof and having a feed duct 14 foradmitting the propellant connected to the upstream end thereof. AT-shaped piston 16 having a piston head 18 and a piston rod 20 iscapable of sliding longitudinally inside the body 10 in practicallysealed manner (e.g. sealed by a sliding piston ring 30), moving betweena rest, first position at the upstream end of the body and an operating,second position downstream therefrom, the piston subdividing the bodyinto a combustion chamber 22 disposed in front of the piston head and inwhich the combustion gases form, and an annular injection chamber 24surrounding the piston rod and in which the propellant is stored priorto injection into the combustion chamber. The piston 16 is guided in thebody 10 firstly by the piston head 18 which is capable of sliding in afirst bore 26 formed in the body 10 parallel to its longitudinal axis,and secondly by the piston rod 20 which is itself capable of sliding ina second bore 28 in coaxial alignment with the preceding bore and formedat an upstream end of the body 10 on the axis of the feed duct 14.

The liquid propellant from a separate source 21 is conveyed, optionallyafter passing through an isolating valve 32 (used to confine thepropellant in sealed manner while it is being stored), to an injectionvalve 34 from which it can be introduced into the annular injectionchamber 24 along the second bore 28 via the feed duct 14. Thereafter itis injected into the combustion chamber 22 via injection channels 36formed through the piston head 18, after which the propellant isdecomposed into a high temperature gas by passing through a catalyticbed 38 placed directly on the front face of the piston head 20 thatfaces into the combustion chamber 22.

The combustion chamber 22 includes a downstream stop 40 that mayadvantageously be constituted by a shoulder of the converging/divergingnozzle 12 and that is designed to limit downstream displacement of theinjector-forming piston 16 (thereby defining the operating secondposition), and the end wall of the injection chamber 24 constitutes anupstream stop 42 for the injector (and thus corresponds to the rest,first position). Return means 44 can also be provided to hold theinjector 16 in the rest, first position occupied by the engine beforebeing fed with propellant (or during storage). It should be observedthat in most cases, friction forces between the moving injector and thechamber body can suffice to hold the injector in the rest firs position.

Where it joins the injection chamber 24, the second bore 28 has a recess46 which serves to define a flow section through which propellant isinjected into said chamber when the injector is in its operatingposition pressed against the downstream stop 20. Naturally, the pistonrod 20 must be dimensioned so as to ensure that said flow section existsin the operating position and, conversely, to ensure that the recess 46is completely closed off in the rest position when the injector piston16 is in upstream abutment against the end wall 42 of the injectionchamber 24.

A coil 48 and a pole piece 50 forming an open torus may be disposed onan upstream outside wall of the body 10 of the rocket engine so thatwhen appropriately excited by control means 52 (not shown), it enablesthe injector to be locked in its rest position bearing against theupstream stop 42, in particular when the pressure in the feed duct 13would otherwise tend to urge said injector towards its operatingposition.

The operation of a miniaturized rocket engine can be described asfollows in simplified manner with reference to FIGS. 2a and 2b. In FIG.2b, the rest position of the piston (upstream position) corresponds to astroke d=-4 mm, and the operating position (downstream position)corresponds to a stroke d=0 mm. Prior to use, the return means 44guarantee that the injector 16 is pressed against the upstream stop 42.On first use, the pressurized propellant coming from the tank 21 flowsinto the feed duct 14 after the injection valve 34 has been opened, thusapplying pressure to the free end of the piston rod 20. This pressurefrom the propellant generates a force that opposes the action of thereturn means 44, and the friction forces of the piston rings 30, inparticular, and also the force on the piston head 18 and on thecatalytic bed 38 as generated by the pressure that exists in thecombustion chamber 22, and as a result the injector is displaced towardsthe downstream stop 40 (curve of FIG. 2b).

During this first displacement, the flow section of the recess 46 isdisengaged and a flow of propellant is then established between the feedduct 14 and the injection chamber 24 via the bore 28 (dashed line curve54 in FIG. 2a represents the integral of this "filling" flow rate, i.e.it shows the mass of propellant contained in the injection chamber as afunction of elapsed time). Once the chamber is full, and after a hammershock, the pressure in the chamber causes the propellant to be injectedinto the combustion chamber 22 via the channels 36. The propellant thenpasses through the catalytic bed 38 where it is decomposed into hightemperature gas. It may nevertheless be observed that given the timerequired for such decomposition, a certain amount of propellant readyfor decomposition can accumulated in the combustion chamber 22. Thedecomposition gives rise to a sudden increase in the pressure inside thecombustion chamber to a value that is greater than the pressurizationpressure of the propellant and sufficient to generate forces that opposethe force causing the piston 16 to advance, as represented by the solidline curve 56 in FIG. 2a, in which there can be seen an "ignition peak"58 which is also typical in certain conventional engines.

The piston 16 is then subjected to second displacement in the oppositedirection to the preceding displacement, from the downstream stop 40towards the upstream stop 42 (the end wall of the injection chamber 24),thereby deliberately giving rise to propellant being injected into thecombustion chamber 22 through the channels 36 and the catalytic bed 38and also returning a portion (a very small portion, and in any event aportion that is negligible compared with the total volume of theinjection chamber 24) back into the feed duct 14 through the bore 28 viathe flow section of the recess 46 which is still open but which isdecreasing very quickly as the piston 16 recoils. Once the flow sectionis completely closed off, the propellant present in the injectionchamber is pressurized even more strongly by a differential pistoneffect and is thus forced to inject into the combustion chamber 22passing via the channels, and the catalytic bed, thereby giving rise toa rapid increase in the chamber pressure (for example to reach a chamberpressure of 650 bars compared with a nominal feed pressure of 20 bars).This pressure emphasizes rearward displacement of the piston 16 until itstabilizes when the piston makes contact with the end wall of theinjection chamber 24 at a value that is determined by the geometricaland functional characteristics of the rocket engine. In this position,the piston has returned to its initial, rest location against theupstream stop 42, but the pressure conditions in the combustion chamber22 are now very different. It is only after a determined length of timethat the combustion gases exhausted via the nozzle 12 enable thepressure in the combustion chamber 22 to decrease rapidly so that itreturns to the initial pressure conditions, after which new cycles canbe engaged without interruption, automatically and one after another,for so long as the feed pressure continues to act on the piston rod 16(i.e. until no more propellant remains in the tank 21).

Naturally, it is also possible to stop operation of the engine at anytime by preventing the piston 16 from moving, either by reducing thefeed pressure (in practice by closing the injection valve 34), or byexciting the coil 50 which is appropriately dimensioned to be able toblock the piston, even if the pressure in the feed duct is at itsnominal value. When blocking is performed in this manner, it isnaturally appropriate for the piston 16 to be made at least in part outof a ferromagnetic material, it being assumed that the upstream end ofthe body 10 is not ferromagnetic.

FIGS. 3 and 3a are diagrammatic longitudinal section views of a rocketengine of the invention fed with two propellants, preferably hypergolic.Although it is advantageous and desirable for the propellants to behypergolic since that gives rise to spontaneous ignition when they areinjected into the combustion chamber, the use of non-hypergolic isnaturally also possible, providing suitable ignition means are providedin the combustion chamber.

In FIG. 3, as in the preceding embodiment having a single propellant,the rocket engine comprises a chamber body 60 having a thrust nozzle 62connected to its downstream end and feed ducts 64a and 64b connected toits upstream end to deliver propellants via an intermediate part 65. Apiston 66 including a piston head 68 capable of sliding in a first boreof cylindrical shape 76 and an annular piston rod 70 itself capable ofsliding in a first annular cavity 78 in coaxial alignment with the firstbore 76 is capable of being displaced longitudinally in practicallysealed manner (see for example the sliding piston ring 30) within thebody 60, between a rest first position and an operating second position,subdividing said body into a combustion chamber 72 in front of thepiston and both a central injection chamber 74a and an annular injectionchamber 74b. The annular piston rod 70 is extended by a cylindricalannular wall 80 projecting rearwards and itself capable of sliding in asecond annular cavity 82 in coaxial alignment with the first annularcavity 78 and formed in an upstream end of the chamber body 60 in theend wall of said first cavity 78. The second annular cavity 82 isconnected to the outside atmosphere via an opening 84 formed in thechamber body 60, e.g. perpendicularly to the longitudinal axis of therocket engine. O-rings 86 provide sealing for the second annular cavity82, thereby closing off the upstream end of the chamber body 60 insealed manner, while defining two independent concentric cavities 88 and90 on opposite sides of the wall 80 for the purpose of injecting thepropellants.

As in the preceding embodiment, the combustion chamber 72 includes adownstream stop 96, e.g. constituted by a shoulder on theconverging/diverging nozzle 62, and designed to limit downstreamdisplacement of the injector-forming piston 66 (thus defining theoperating, second position). The rearwards limit on displacement of thepiston 66 (corresponding to its rest, first position) is preferablydetermined by an upstream abutment 96 constituted by the end wall of thecentral injection chamber 74a, or 74b, but the end wall of the firstannular cavity 78 or the end wall of the second annular cavity 82 couldalso perform this function of upstream stop.

Where it joins the injection chambers 74a and 74b, the first annularcavity 78 includes a recess 100 which serves to define a flow sectionfor injecting the propellants into said chambers when the injector is inits operating position in abutment against the downstream stop 96. It isclear that the piston rod 70 which slides in said cavity 78 must bedimensioned so as to ensure that said flow section exists in theoperating position, and conversely, to ensure that the recess 100 iscompletely shut off in the rest position, with the injector piston 66then being in upstream abutment 98 against the end wall of the centralinjection chamber 74a.

This configuration makes it possible to use a pulse rocket engine thatis optimized with respect to the leakage path between the independentannular cavities 88 and 90, and the cavity 82 which is connected to the"atmosphere" via the opening 84, since the pressure in the cavities 88and 90 is substantially identical to that in the feed ducts, and is thusat low pressure. This feature is taken advantage of in a compact variantdescribed in further detail with reference to FIG. 4.

The hypergolic propellants coming from separate sources (not shown) areintroduced into the concentric cavities 88 and 90 from the feed ducts14a and 14b via passages 102, 104 formed in the upstream end of the body60 and in the part 65 and leading into the first annular cavity 78, andthey are injected into the combustion chamber 72 through injectionchannels 92 and 94 pierced through the moving injector 66 andadvantageously disposed in the form of a ring of channel pairs.

The operation of the two-propellant pulse rocket engine is similar tothat of the engine of FIG. 1, but it should be observed that since thepropellants are hypergolic, the ignition means needed with a singlepropellant (e.g. a catalytic bed) are no longer necessary. Naturally,ignition means would be essential if the propellants werenon-hypergolic. It may also be observed that in order to avoidovercrowding the drawing, means for stopping the engine (valve, magneticdevice, or other analogous device) by locking the piston in its restposition are not shown.

FIG. 3a shows a pulse rocket engine that is optimized with respect toleakage paths between the central and annular injection chambers 74a and74b which are subjected to intense pressure, and the feed ducts 64a and64b which are pressurized only slightly, since during operation (goingfrom the operating position towards the rest position), flow through theflow sections 100a and 100b takes place in the opposite direction tothat in which the piston 66, 70 is moving. This gives rise to leakagebeing braked and thus to the amount of said leakage being reduced.Advantage can be taken thereof by relaxing manufacturing tolerances.

FIG. 4 shows a compact variant of the FIG. 3 rocket engine. In thisvariant, the cylindrical annular wall 80 for separating the propellantswhich is subjected to modest pressure forces only is replaced by atleast one resilient bellows 110 which is fixed, preferably by welding toensure maximum sealing, firstly to the annular rod of the piston 70 andsecondly to the upstream end of the intermediate part 65. In addition,in order to avoid any risk of jamming during displacement of theinjector 66, the piston rod 70 which moves externally in a bore 112 ofthe chamber body slides internally on a floating part 114 which acts asa plunger. This part, which is nevertheless installed to tighttolerances, is secured in flexible manner to the upstream end of thechamber body by resilient fixing means of the resilient pin type 116,for example, so as to make the connection between the injector and thebody leakproof and insensitive to jamming.

FIGS. 5 to 8 are highly schematic theoretical diagrams showing variantembodiments of the pulse rocket engine of the invention, for one or twopropellants.

In the versions of FIGS. 5 and 6, there can be seen two embodiments ofmeans for locking the injector piston 16 in its initial, rest positionfirstly by means of an electrohydraulic unit 120 and secondly by meansof an electromechanical unit 130. In FIG. 5, the piston rod 20 has acavity 122 capable of sliding on the inside piston 124 and its other endopens out into the feed duct 14 where it can be closed by a solenoidvalve 128. In the example shown, it will be observed that ignition means125 are provided in the combustion chamber 22. The above-specifiedlocking means operate in a manner similar to that of hydraulicactuators, with propellant admission via the valve 128 and the bore 126into the cavity 122 serving, under the effect of pressure forces actingon the various members of the device after the solenoid valve has beenclosed, to lock the piston rod 20 in a determined position (and inparticular in its rest position). It may be observed that these lockingmeans have the advantage of making it very easy to obtain pulses ofdifferent shapes merely by acting on the solenoid valve 128 in such amanner as to control the recoil of the injector 16. In FIG. 6, theannular wall 80 separating the propellants includes towards its rear enda projection which is pierced by an orifice 132 suitable for receiving afinger 134 actuated by an electromagnetic 136 and held in position by areturn spring 138. The injector piston 66 is locked merely by excitingthe electromagnet 136 which urges the finger 134 into the orifice 132.The locking finger preferably slides perpendicularly to the displacementdirection of the piston rod 70 (and thus of the wall 80) since undercircumstances, forces are minimized.

FIGS. 7a and 7b show embodiment details of the injector 66. In FIG. 7a,propellant injection into the combustion chamber 72 no longer takesplace via two channels pierced through the injector but takes placesimultaneously via a central orifice 140 of the injector and via a gap142 between the periphery of the injector 66 and the bore 76 in thechamber body 60, which gap may merely be the determined toleranceclearance that exists between the chamber body and the injector. In FIG.7b, the front face of the injector 66 facing the combustion chamber 72includes a forwardly projecting portion 144 designed to close the throatof the thrust nozzle to a greater or lesser extent. This portion 144serves above all to optimize the operation of the rocket engine, but itmay also act as a downstream stop.

Finally, FIG. 8 shows a variant embodiment of a pulse rocket engine fedwith two propellants. In this embodiment, the injector 66 is providedwith a piston head 68 which also constitutes an injection wall for thecombustion chamber 72 whose chamber body 160 is connected to a thrustnozzle 162. This moving assembly comprising the injector plus thechamber body plus the nozzle can move longitudinally in the body of therocket engine 60, a downstream stop 196 constituted by an inwardlydirected flange on the body 60 limits such longitudinal displacement.Naturally, as in each of the preceding embodiments, the piston rod isdimensioned in such a way as to enable it to define a flow section forthe propellants when said engine is in its operating position with theinjector pressed against the downstream stop 196. As shown in thefigure, this flow section can be the result merely of relativedisplacement between the piston rod and the bore for said rod, withoutit being necessary to form any hollow in said bore, like the embodimentsshown in FIGS. 1, 3, and 4.

The person skilled in the art can naturally envisage other variants andadaptations based on the embodiments described above, without goingbeyond the ambit of the invention. Thus, for example, it is notessential for the chamber body to have a section that is circular, andit is entirely possible to envisage a non-circular shape for the slidingpiston together with a nozzle throat section that is substantiallyrectangular.

Because it is extremely compact, the present invention is particularlysuited to miniaturized thrust systems. In satellites, it is particularlyappropriate for use in auxiliary or emergency systems. Such systems canalso be selected for aerodynamic maneuvering of missiles as areplacement for conventional control surfaces. More generally, in thefield of aviation, such pulse rocket engines may constitute testmachines for generating vibrations for studying the mechanical behaviorof very large structures. It should also be observed that there is apossible application to anti-skidding systems for motor driven landvehicles, where the variable thrust with a short response time asprovided by such a rocket engine can make it possible to compensate theloss of grip caused by centrifugal force.

Finally, it may be observed that although rocket engines of theinvention provide intermittent thrust, it is possible to combine aplurality of engines in a thrust unit so as to obtain continuous thrustlike that of a piston engine.

We claim:
 1. A liquid propellant rocket engine comprising a chamber bodyconnected to a converging/diverging thrust nozzle and to at least onelightly pressurized propellant feed duct, said body including a bodybore in which an injector-forming piston can move longitudinally betweena rest first position and an operating second position, the pistoncomprising a piston head and a piston rod, and subdividing the chamberbody bore into a combustion chamber in front of the piston, and at leastone injection chamber surrounding the piston rod, wherein the chamberbody further includes a flow section that is closed in said restposition to prevent any propellant being fed from the injectionchamber(s) and that is opened when the piston moves from said restposition to said operating position in order to allow the injectionchamber(s) to be filled via said flow section directly from the feedduct(s), the propellant(s) being injected into the combustion chambervia injection channels formed between the combustion chamber and theinjection chamber(s), displacement of the piston being a function, inparticular, of the pressures in the feed duct(s), in the injectionchamber(s), and in the combustion chamber.
 2. A rocket engine accordingto claim 1, wherein said flow section is provided by a recess formed atthe inlet to the injection chamber(s) in the body bore along which thepiston rod slides.
 3. A rocket engine according to claim 1, wherein saidpropellant injection channels into the combustion chamber areconstituted by clearance existing between the bore of he chamber bodyand the peripheral surface of the piston head.
 4. A rocket engineaccording to claim 1, wherein said piston head includes a forwardlyprojecting portion on its front face facing into the combustion chamberwhich portion can close the throat of the thrust nozzle in full or inpart during displacement thereof from the rest position to the operatingposition.
 5. A rocket engine according to claim 1, further including adownstream stop formed by a shoulder of the thrust nozzle and designedto limit displacement of the piston in said operating position.
 6. Arocket engine according to claim 1, fed with two propellants, whereinthe piston rod is annular in shape and defines both a central injectionchamber and an annular injection chamber, said piston rod including awall that projects rearwards and slides in a second annular cavity thatis in coaxial alignment with a first annular cavity in which the pistonrod slides, such that such wall defines two concentric cavities forfeeding the two propellants from their respective feed ducts.
 7. Arocket engine according to claim 6, wherein said wall is replaced by atleast one resilient bellows fixed, preferably by welding, to an upstreamend of the chamber body.
 8. A rocket engine according to claim 6,wherein the piston rod slides both externally in a bore of the chamberbody and internally on a plunger-forming part fixed to the upstream endof said body by resilient fixing means, such that the connection betweenthe piston and the body provides good sealing and limits any risk of thepiston jamming during displacement thereof.
 9. A rocket engine accordingto claim 1, fed with a single propellant, the engine including ignitionmeans for igniting combustion.
 10. A rocket engine according to claim 9,wherein said ignition means comprise a catalytic bed disposed on thefront face of the piston head facing towards the combustion chamber. 11.A rocket engine according to claim 1, also including return means forholding the piston in the rest, first position, in particular in theabsence of any propellant feed.
 12. A rocket engine according to claim1, wherein, to cause it to stop operating, it further includes lockingmeans for locking the piston in the rest, first position.
 13. A rocketengine according to claim 12, in which at least a portion of the piston,and a piston rod, is made of a ferromagnetic material, and at least theupstream end of the chamber body is made of a nonferromagnetic material,wherein said locking means are constituted by an electromagnetic unitwhich, when excited by control means, causes displacement of the pistonto stop.
 14. A rocket engine according to claim 12, wherein said lockingmeans are constituted by an electromagnetic unit comprising anelectromagnet which, when excited, causes a locking finger to move intoan orifice formed in the piston rod or in any other part securedthereto.
 15. A rocket engine according to claim 12, wherein said lockingmeans are constituted by an electrohydraulic unit constituted by anactuator fed from at least one feed duct via a solenoid valve, and onwhich a cavity formed in the piston rod can slide in such a manner thatsaid piston rod is capable of being locked in a determined positionunder the effect of the pressure forces involved.
 16. A rocket engineaccording to claim 1, having noncircular sliding geometry for the pistonand a nozzle throat that is of substantially rectangular section.